Combustion systems

ABSTRACT

A combustor for a gas turbine engine includes an inner combustor wall and an outer combustor wall radially outboard of the inner combustor wall. The inner and outer combustor walls define a combustion space therebetween with an upstream inlet and a downstream outlet. A combustor dome connects between inner and outer combustor walls at the upstream inlet of the combustion space. The combustor dome includes a plurality of circumferentially spaced apart nozzles and a plurality of tiles mounted to the combustor dome to fluidly separate a downstream side of the combustor dome from an upstream side of the combustor dome.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application in a continuation of U.S. patent application Ser. No.14/749,275 filed Jun. 24, 2015 which is incorporated by reference hereinin its entirety.

BACKGROUND OF THE INVENTION 1. Field of the Invention

The present disclosure relates to combustion, and more particularly tocombustion systems such as used in gas turbine engines.

2. Description of Related Art

A variety of devices and methods are known in the art for combustion. Ofsuch devices, many are directed to combustion for gas turbine engines.Traditionally, fuel in gas turbine engines is supplied through fuelinjectors for combustion within a combustor. The fuel injectors wereconnected to the upstream wall, or combustor dome, of the combustor andwere required to be small enough to be removed without opening theengine case completely. This allowed for frequent changing out ofinjectors. But more and more demanding performance requirements aredriving an increasing trend towards injectors with larger and largernozzles.

Such conventional methods and systems have generally been consideredsatisfactory for their intended purpose. However, there is still a needin the art for improved combustion systems. The present disclosureprovides a solution for this need.

SUMMARY OF THE INVENTION

A combustion system including a combustor dome including a plurality ofcircumferentially spaced apart nozzles. The system also includes aplurality of tiles mounted to the combustor dome to fluidly separate adownstream side of the combustor dome from an upstream side of thecombustor dome.

Each tile can be mounted around a respective one of the nozzles. Eachtile can be spaced apart from the respective one of the nozzles in anaxial direction between the upstream and downstream sides of thecombustor dome to define an outer air circuit between the tile and therespective one of the nozzles. Each nozzle can include a downstreaminsulation shell with a standoff extending in a downstream directionfrom the nozzle, wherein the standoff spaces the tile from the nozzle tomaintain space for the outer air circuit. Each tile can include a radialslot that receives the standoff. Each tile can define a convergingfrustoconical surface that defines a portion of the outer air circuit.

Each nozzle can include a radial swirler on the upstream side of thecombustor dome in fluid communication with a nozzle outlet on thedownstream side of the combustor dome, wherein an inner air circuit isdefined from the radial swirler to the nozzle outlet. The inner andouter air circuits can be configured to issue 40% to 50% of airflowpassing through the combustor dome through the inner air circuit, and toissue 50% to 60% of the airflow through the outer air circuit, whereinthe total airflow of the inner and outer airflow circuits does notexceed 100%. Struts can support each nozzle in the combustor dome andobstruct flow only to the outer air circuit wherein the outer aircircuit is unobstructed by the struts. A channel forming the inlet ofthe outer air circuit can be proximate an upstream surface of the tilefor cooling of the tile by airflow passing into the radial swirler

Each tile can include a radial edge with a radially extending groovedefined therein, wherein each tile includes a radial edge with aradially extending tongue thereon, and wherein each tile engagescircumferentially adjacent tiles wherein the tongues are engaged in thegrooves. Each tile can include a ceramic matrix composite (CMC)material.

Each tile can be confined axially with the combustor dome. The tiles canbe confined axially by an outer bayonet ring proximate a radially outerperimeter defined by the tiles, and by an inner bayonet ring proximate aradially inner perimeter defined by the tiles, wherein each of thebayonet rings is interlocked with bayonet features in respective innerand outer rings of the combustor dome.

A combustor for a gas turbine engine includes an inner combustor walland an outer combustor wall radially outboard of the inner combustorwall. The inner and outer combustor walls define a combustion spacetherebetween with an upstream inlet and a downstream outlet. A combustordome as described above connects between inner and outer combustor wallsat the upstream inlet of the combustion space.

The tiles can be confined axially by an outer bayonet ring proximate aradially outer perimeter defined by the tiles, and by an inner bayonetring proximate a radially inner perimeter defined by the tiles, whereineach of the bayonet rings is interlocked with bayonet features inrespective inner and outer rings of the combustor dome, wherein theouter bayonet ring slidingly supports a cold side combustor liner of theouter combustor wall, and wherein the inner bayonet ring slidinglysupports a cold side combustor liner of the inner combustor wall. It isalso contemplated that the outer bayonet ring can form a cold sidecombustor liner of the outer combustor wall, and the inner bayonet ringcan form a cold side combustor liner of the inner combustor wall.

These and other features of the systems and methods of the subjectdisclosure will become more readily apparent to those skilled in the artfrom the following detailed description of the preferred embodimentstaken in conjunction with the drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

So that those skilled in the art to which the subject disclosureappertains will readily understand how to make and use the devices andmethods of the subject disclosure without undue experimentation,preferred embodiments thereof will be described in detail herein belowwith reference to certain figures, wherein:

FIG. 1 is a cross-sectional side elevation view of an exemplaryembodiment of a combustion system constructed in accordance with thepresent disclosure, showing a combustor in fluid communication between acompressor and a turbine in a gas turbine engine;

FIG. 2 is a perspective view of a portion of the combustion system ofFIG. 1, showing a combustor dome with integrated fuel manifold andnozzle components;

FIG. 3 is an axial outlet end view of the combustor dome of FIG. 2;

FIG. 4 is a perspective view of a portion of the combustion system ofFIG. 1, showing standoffs for spacing a tile of the combustor dome fromone of the nozzles;

FIG. 5 is a perspective view of the portion of the combustion system ofFIG. 4, showing the tile assembled into the frame of the combustor dome;

FIG. 6 is a radial cross-sectional view of the portion of the combustionsystem of FIG. 5, showing the inner and outer air circuits;

FIG. 7 is a perspective view of the portion of the combustion system ofFIG. 5, schematically showing air flow into the radial swirler coolingthe upstream side of the tile;

FIG. 8 is an axial elevation view of the tile of FIG. 5, showing thedownstream or hot side of the tile;

FIG. 9 is an axial elevation view of the tile of FIG. 8, showing theupstream or cold side of the tile;

FIG. 10 is a perspective view of the tile of FIG. 8, showing the tongueand groove for circumferential assembly with other tiles;

FIG. 11 is a perspective view of the portion of the combustion system ofFIG. 7, schematically indicating engagement of the bayonet rings;

FIG. 12 is a perspective view of a portion of the combustion system ofFIG. 11, showing the bayonet features; and

FIG. 13 is a radial cross-sectional elevation view of the portion of thecombustion system of FIG. 11, showing the engagement of the bayonetrings with the inner and outer combustor walls.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

Reference will now be made to the drawings wherein like referencenumerals identify similar structural features or aspects of the subjectdisclosure. For purposes of explanation and illustration, and notlimitation, a partial view of an exemplary embodiment of a combustionsystem in accordance with the disclosure is shown in FIG. 1 and isdesignated generally by reference character 100. Other embodiments ofcombustion systems in accordance with the disclosure, or aspectsthereof, are provided in FIGS. 2-13, as will be described. The systemsand methods described herein can be used for combustion in gas turbineengines.

Combustion system 100 for a gas turbine engine includes a combustor dome102 connecting between the inner and outer combustor walls 104 and 106,respectively, to form a combustor. Outer combustor wall 106 is radiallyoutboard of inner combustor wall 104, such that the inner and outercombustor walls 104 and 106 define an annular combustion space 108therebetween with an upstream inlet 110 at combustor dome 102, e.g., forreceiving fuel and compressed air, and a downstream outlet 112, e.g.,for providing pressurized combustion products to a downstream turbine114. Compressor 116 is connected in fluid communication with combustordome 102 to provide compressor discharge air to the inlet 110 of thecombustion space 108. Compressor 116, combustor dome 102, and turbine114 are annular components centered on engine centerline axis A.

Referring now to FIG. 2, combustor dome 102 includes a frame 101 with anintegral fuel manifold 118 having an inlet 120 and nozzle components 122of a plurality of nozzles circumferentially spaced around combustor dome102. Fuel manifold 118 and nozzle components 122 are integrated with thecombustor dome 102 for fluid communication from inlet 120 to the nozzlecomponents 122. Those skilled in the art will readily appreciate thatwith the fuel manifold 118 and nozzle components 122 integrated into thecombustor dome 102, the traditional nozzle feed arms extending fromoutside the combustor and extensive fuel manifolding external to thecombustor for feeding each nozzle are eliminated.

Inlet 120 is radially outboard of the nozzle components 122 forsupplying fuel to the manifold from a source external to the combustordome. A flexible fitting 124 can be used to connect inlet 120 with thefuel source, e.g. by way of a fuel control system. If required for agiven application, multiple fittings 124 and inlets 120 can be used.Radially inserted slip pins 126 can be used to support combustor dome102 from engine case 128, so as to accommodate differential thermalexpansion and contraction between combustor dome 102 and the surroundingcomponents such as engine case 128 (shown in FIG. 1) and the fuelsystems connected by fitting 124.

With continued reference to FIGS. 2 and 3, frame 101 includes inner andouter combustor dome rings 132 and 130. Fuel manifold 118 can form theouter combustor dome ring 130, wherein the nozzle components 122 extendinward from the outer combustor dome ring radially inward to an innercombustor dome ring 132. The inner combustor dome ring 132 iscircumferentially segmented, with one segment per nozzle, to accommodatethermal expansion and contraction of the combustor dome. Another way ofaccommodating thermal expansion and contraction in combustor dome 102 inaddition to or in lieu of segmenting inner ring 132 is to segment theentire combustor dome, e.g., along the dashed lines in FIG. 3, whereineach segment would have its own inlet 120 and separate manifold 118 forfeeding its subset of nozzles. The structure of combustor dome 102 asshown in FIGS. 2 and 3 can be fabricated and brazed, additivelymanufactured, e.g., complete with internal nozzle details, or can beproduced using any other suitable processes.

With reference now to FIG. 4, a portion of combustion system 100 isshown tile 134 removed (tile 134 can be seen in FIG. 1). Combustionsystem 100 includes a plurality of tiles 134 mounted to the frame 101 ofcombustor dome 102 to fluidly separate a downstream side of thecombustor dome 102 from an upstream side of the combustor dome 102.Nozzle components 122 of each nozzle include a downstream insulationshell 136 with standoffs 138 extending in a downstream direction D fromthe nozzle, i.e., the axial direction between the upstream anddownstream sides of the combustor dome 102. When the tile 134 isassembled onto combustor dome 102, as shown in FIG. 5, the standoffs 138spaces the tile 134 apart from the nozzle to maintain space defining anouter air circuit 140 of the nozzle between tile 134 and shell 136 ofthe nozzle, as indicated in FIG. 6.

With continued reference to FIG. 6, each tile 134 is mounted around arespective one of the nozzles. The nozzle components 122 of each nozzleinclude a radial swirler 142 on the upstream side of combustor dome 102in fluid communication with a nozzle outlet 144 on the downstream sideof combustor dome 102. An inner air circuit 146 is defined from radialswirler 142 to nozzle outlet 144. Inner and outer air circuits 142 and140 are configured to issue 40% to 50% of airflow passing throughcombustor dome 102 through the inner air circuit 146, and to issue 50%to 60% of the airflow through the outer air circuit 140, wherein thetotal airflow of the inner and outer airflow circuits does not exceed100%. Substantially all of the air supplied to the combustor space 108enters by way of the nozzles of combustor dome 102. Those skilled in theart will readily appreciate that the air flow splits above can beadjusted as required on an application by application basis withoutdeparting from the scope of this disclosure.

Struts 148, also shown in FIGS. 2 and 4, support each nozzle in theframe 101 of combustor dome 102 and only modestly obstruct flow, e.g.,air discharged from compressor 116, only to the outer air circuit 140.The outer air circuit 140 and radial swirler 142 are unobstructed bystruts 148. As viewed from the upstream side in FIG. 7, the radialchannel 141 labeled in FIGS. 4-6, forming the inlet for outer aircircuit 140 is located intimate and proximate the upstream surface oftile 134 for cooling of the tile 134 by airflow passing into the outerair circuit 140. The proximity of the channel 141 of outer air circuit140 ensures high flow along the upstream surface of tile 134 and thislarge quantity of air flowing radially into the nozzle scrubs the backside of the tile 134, cooling it from the back side before the airenters the combustion volume 108. This cooling flow is indicatedschematically by the converging arrows in FIG. 7. The back side wall 143of radial swirler 142 can be made of a composite material, or any othersuitable material.

With reference now to FIG. 8, tile 134 has a segment-shaped profile witha radially outer perimeter 150 and opposed radially inner perimeter 152.The downstream surface of tile 134 that is shown in FIG. 8 defines acircular outlet 144 for the outer air circuit 140, so outlet 144 isdimensioned to conform to the requirement for nozzle air flow. As shownin FIG. 9, tile 134 defines a converging frustoconical surface 154 thatdefines a portion of the outer air circuit 140 shown in FIG. 6. Tile 134is thus integrated into the nozzle as one of the nozzle components 122.Tile 134 includes radial slots 156 that receive the standoff 138 thatare shown in FIG. 4. Slots 156 maintain concentricity between the nozzleand tile 134.

As shown in FIG. 10, each tile 134 include a radial edge with a radiallyextending groove 158 defined therein. Each tile 134 also includes aradial edge with a radially extending tongue 160 thereon. With all ofthe tiles 134 engaged around the circumference of combustor dome 102,each tile 134 engages circumferentially adjacent tiles 134 wherein thetongues 160 are engaged in the grooves 158. The tongue and groovearrangement allows circumferential positional adjustment of tiles 134,reducing or eliminating radial growth arising from thermal expansion,while maintaining a seal between the adjacent tiles 134.

Referring now to FIG. 11, each tile 134 is confined axially with thecombustor dome 102, i.e. against movement in the direction D indicatedin FIG. 6. The tiles 134 are confined axially by an outer bayonet ring162 proximate the radially outer perimeter 150 defined by the tiles 134,and by an inner bayonet ring 164 proximate the radially inner perimeter152 defined by the tiles 134. This allows tiles 134 to contract andexpand radially under thermal cycles. Each of the bayonet rings 162 and164 is a continuous ring, however it is also contemplated that splitrings, or multiple ring segments could be used for one or both. Each ofthe bayonet rings 162 and 164 is interlocked with bayonet features inrespective inner and outer rings 132 and 130 of the combustor dome 102so that bayonet rings 162 and 164 can be engaged and disengaged fromcombustor dome 102 by rotation in the direction indicated by the doublearrows in FIG. 11. FIG. 12 shows a bayonet feature 166 of the outer ring130 of frame 101 interlocked with a bayonet feature 168 of outer bayonetring 162. The bayonet feature 168 of outer bayonet ring 162 can beintroduced into and removed from outer ring 130 through notch 170defined in outer ring 130. Inner bayonet ring 164 and inner ring 132 offrame 101 have a similar engagement with bayonet features 174 and 176,and notch 172, which are shown in FIG. 11. Each nozzle has acorresponding arrangement of bayonet features 166, 168, 174, and 176,however it is also contemplated that some of the bayonet features can beeliminated as long as enough are included for secure tiles 134 in place.

With reference now to FIG. 13, outer bayonet ring 162 slidingly supportscold side combustor liner 178 of the outer combustor wall 106. Innerbayonet ring 164 slidingly supports cold side combustor liner 180 of theinner combustor wall 104. Those skilled in the art will readilyappreciate that optionally the outer bayonet ring 162 can form a coldside combustor liner of the outer combustor wall 106, e.g., by makingouter bayonet ring 162 and liner 178 integral, and that inner bayonetring 164 can form a cold side combustor liner of the inner combustorwall 104, e.g., by making inner bayonet ring 164 integral and liner 180integral. Bayonet rings 162 and 164 can be made of any suitable materialincluding metallic or composite materials.

Any suitable material can be used for tiles 134, such as SiC, SiCceramic matrix composite (CMC) material. SiC and SiC CMC materials arelight weight and capable of withstanding very high temperatures typicalof combustion without necessarily requiring active cooling. Thesematerials conduct heat relatively well, so backside cooling as describedabove can be sufficient to protect against direct combustion heating.Adequate air flow through the nozzle provides sufficient backsidecooling. These materials also simplify design due to their lowcoefficient of thermal expansion. Using a design with partitioned tilesas described herein helps decouple thermal growth differentials betweenthe outer ring of the combustor dome and the tiles. The components ofcombustor dome arrangements described herein tend to be loaded in theaxial direction, pushing the components together axially. Compositematerials perform favorably under such loading. This coupled with thelow thermal growth, which can beneficially match the low thermal growthof components conducting cool fuel therethrough, such as the fuelmanifold 118, make composites a good material for elements such asradial swirler backside wall 143, tiles 134, and bayonet rings 162 and164. Elements conducting fuel therethrough, such as frame 101 ofcombustor dome 102, can advantageously be metallic.

The nozzles and combustor domes described herein provide many potentialadvantages over the traditional configurations. The fuel nozzles can nowintegrate a large majority of the combustor and fuel manifold into thenozzle structure. The nozzles and combustor dome can be engineered formaintenance/replacement to coincide with the replacement of other enginecomponents such as turbines, where the engine case must be opened.Therefore, the nozzles do not need to be small enough to be extractedthrough the case and combustor. The larger nozzles can provideadvantages including more efficient mixing large amounts of fuel andair, and the nozzle air can be used to cool the combustor, simplifyingcombustor design. The combustor itself can be substantially simplifiedwith combustor domes such as disclosed herein. Thermal mismatch betweenthe combustor and engine case can be decoupled from the nozzle/combustorinterface. The weight of heavy stems (feed arms or the like), flanges,and engine case reinforcement can be reduced or eliminated. Having thefuel manifold inside the engine, i.e. integrated in the combustor dome,means the fuel manifold resides closer to the nozzles, reducing theamount of wetted walls of the fuel system. Common features between thenozzles and combustor can be unitized for construction.

The methods and systems of the present disclosure, as described aboveand shown in the drawings, provide for combustion systems with superiorproperties including improved combustion efficiency, reduced weight, andreduction of fuel lines. While the apparatus and methods of the subjectdisclosure have been shown and described with reference to preferredembodiments, those skilled in the art will readily appreciate thatchanges and/or modifications may be made thereto without departing fromthe scope of the subject disclosure.

What is claimed is:
 1. A combustion system comprising: a combustor domeincluding a plurality of circumferentially spaced apart nozzles; aplurality of tiles mounted to the combustor dome to fluidly separate adownstream side of the combustor dome from an upstream side of thecombustor dome, wherein the tiles are confined axially by an outerbayonet ring proximate a radially outer perimeter defined by the tiles,and by an inner bayonet ring proximate a radially inner perimeterdefined by the tiles, wherein each of the bayonet rings is interlockedwith bayonet features in respective inner and outer rings of thecombustor dome.
 2. A combustion system as recited in claim 1, whereineach tile is mounted around a respective one of the nozzles.
 3. Acombustion system as recited in claim 2, wherein each tile is spacedapart from the respective one of the nozzles in an axial directionbetween the upstream and downstream sides of the combustor dome todefine an outer air circuit between the tile and the respective one ofthe nozzles.
 4. A combustion system as recited in claim 3, wherein eachnozzle includes a downstream insulation shell with a standoff extendingin a downstream direction from the nozzle, wherein the standoff spacesthe tile from the nozzle to maintain space for the outer air circuit. 5.A combustion system as recited in claim 4, wherein each tile includes aradial slot that receives the standoff.
 6. A combustion system asrecited in claim 3, wherein each tile defines a converging frustoconicalsurface that defines a portion of the outer air circuit.
 7. A combustionsystem as recited in claim 3, wherein each nozzle includes a radialswirler on the upstream side of the combustor dome in fluidcommunication with a nozzle outlet on the downstream side of thecombustor dome, wherein an inner air circuit is defined from the radialswirler to the nozzle outlet, wherein the inner and outer air circuitsare configured to issue 40% to 50% of airflow passing through thecombustor dome through the inner air circuit, and to issue 50% to 60% ofthe airflow through the outer air circuit, wherein the total airflow ofthe inner and outer airflow circuits does not exceed 100%.
 8. Acombustion system as recited in claim 7, wherein struts supporting eachnozzle in the combustor dome modestly obstruct flow only to the outerair circuit and the outer air circuit is unobstructed by the struts. 9.A combustion system as recited in claim 1, wherein each tile is mountedaround a respective one of the nozzles, wherein each tile is spacedapart from the respective one of the nozzles in an axial directionbetween the upstream and downstream sides of the combustor dome todefine an outer air circuit between the tile and the respective one ofthe nozzles, and wherein a channel forming the inlet of the outer aircircuit is proximate an upstream surface of the tile for cooling of thetile by airflow passing into the radial swirler.
 10. A combustion systemas recited in claim 1, wherein each tile includes a radial edge with aradially extending groove defined therein, wherein each tile includes aradial edge with a radially extending tongue thereon, and wherein eachtile engages circumferentially adjacent neighbor tiles wherein thetongues are engaged in the grooves.
 11. A combustor system as recited inclaim 1, wherein each tile includes a ceramic matrix composite (CMC)material.
 12. A combustor for a gas turbine engine comprising: an innercombustor wall; an outer combustor wall radially outboard of the innercombustor wall, wherein the inner and outer combustor walls define acombustion space therebetween with an upstream inlet and a downstreamoutlet; and and a combustor dome connecting between inner and outercombustor walls at the upstream inlet of the combustion space, thecombustor dome including a plurality of circumferentially spaced apartnozzles and a plurality of tiles mounted to the combustor dome tofluidly separate a downstream side of the combustor dome from anupstream side of the combustor dome, wherein the tiles are confinedaxially by an outer bayonet ring proximate a radially outer perimeterdefined by the tiles, and by an inner bayonet ring proximate a radiallyinner perimeter defined by the tiles, wherein each of the bayonet ringsis interlocked with bayonet features in respective inner and outer ringsof the combustor dome.
 13. A combustor as recited in claim 12, whereinthe outer bayonet ring slidingly supports a cold side combustor liner ofthe outer combustor wall, and wherein the inner bayonet ring slidinglysupports a cold side combustor liner of the inner combustor wall.
 14. Acombustor as recited in claim 12, wherein the outer bayonet ring forms acold side combustor liner of the outer combustor wall, and wherein theinner bayonet ring forms a cold side combustor liner of the innercombustor wall.